Gas turbine engine component platform cooling

ABSTRACT

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform having an outer surface and an inner surface that axially extend between a leading edge portion and a trailing edge portion. At least one augmentation feature is disposed on at least one of the leading edge portion and the trailing edge portion of the outer surface of the platform.

This invention was made with government support under Contract No.F33615-03-D-2354-0017 awarded by the United States Air Force andContract No. N00421-99-C-1270-0011 awarded by the United States Navy.

BACKGROUND

This disclosure relates generally to a gas turbine engine, and moreparticularly to a component that can be incorporated into a gas turbineengine. The component can include platform cooling augmentation featuresfor cooling the platform of the component.

Gas turbine engines typically include a compressor section, a combustorsection and a turbine section. During operation, air is pressurized inthe compressor section and is mixed with fuel and burned in thecombustor section to generate hot combustion gases. The hot combustiongases are communicated through the turbine section, which extractsenergy from the hot combustion gases to power the compressor section andother gas turbine engine loads.

Both the compressor and turbine sections of the gas turbine engine mayinclude alternating rows of rotating blades and stationary vanes thatextend into the core flow path of the gas turbine engine. For example,in the turbine section, turbine blades rotate to extract energy from thehot combustion gases that are communicated along the core flow path ofthe gas turbine engine. The turbine vanes prepare the airflow for thenext set of blades.

Blades and vanes are examples of components that may need cooled inorder to withstand the relatively high temperature of the hot combustiongases that are communicated along the core flow path. Typically, coolingis achieved by communicating a dedicated cooling airflow to selectportions of the components.

SUMMARY

A component for a gas turbine engine according to an exemplary aspect ofthe present disclosure includes, among other things, a platform havingan outer surface and an inner surface that axially extend between aleading edge portion and a trailing edge portion. At least oneaugmentation feature is disposed on at least one of the leading edgeportion and the trailing edge portion of the outer surface of theplatform.

In a further non-limiting embodiment of the foregoing gas turbineengine, the platform is an inner diameter platform.

In a further non-limiting embodiment of either of the foregoing gasturbine engines, the component is a turbine vane.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the outer surface is a non-gas path side of the platform andthe inner surface is a gas path side of the platform.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, an airfoil extends from the inner surface of the platform.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the at least one augmentation feature includes a plurality oftrip strips.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the plurality of trip strips are disposed at the trailing edgeportion of the platform.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, each of the plurality of trip strips are angled relative toopposing mate faces of the platform.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, each of the plurality of trip strips include a first portionand a second portion that is transverse to the first portion.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the first portions are angled at a first angle relative to amate face of the platform and the second portions are angled at a secondangle different from the first angle relative to the mate face.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the at least one augmentation feature is disposed at theleading edge portion of the platform.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the at least one augmentation feature is disposed on a portionof the outer surface of the platform that axially overlaps a neighboringcomponent of the gas turbine engine.

A gas turbine engine, according to an exemplary aspect of the presentdisclosure includes, among other things, a compressor section, acombustor section in fluid communication with the compressor section,and a turbine section in fluid communication with the combustor section.At least one of the compressor section and the turbine section includesa first component having a platform that includes an outer surface andan inner surface that axially extend between a leading edge portion anda trailing edge portion and a second component mounted adjacent to thefirst component and including a platform. A portion of the platform ofthe first component axially overlaps a portion of the platform of thesecond component, and the portion of the first platform includes atleast one augmentation feature disposed on the outer surface of theplatform.

In a further non-limiting embodiment of the foregoing gas turbineengine, the first component is a vane and the second component is ablade.

In a further non-limiting embodiment of either of the foregoing gasturbine engines, the at least one augmentation feature includes aplurality of trip strips.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the at least one augmentation feature is disposed on at leastone of the leading edge portion and the trailing edge portion of theplatform of the first component.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the at least one augmentation feature is angled relative to amate face of the platform.

A method of cooling a component of a gas turbine engine according toanother exemplary aspect of the present disclosure includes, among otherthings, cooling a platform of the component with a leakage airflow thatis communicated from a cavity positioned radially inwardly from thecomponent by circulating the leakage airflow over at least oneaugmentation feature that is disposed on an outer surface of theplatform.

In a further non-limiting embodiment of the foregoing method of coolinga component of a gas turbine engine, the leakage airflow is not adedicated cooling airflow that is communicated inside of the component.

In a further non-limiting embodiment of either of the foregoing methodsof cooling a component of a gas turbine engine, the at least oneaugmentation feature is disposed on a trailing edge portion of theplatform.

The various features and advantages of this disclosure will becomeapparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a schematic, cross-sectional view of a gas turbineengine.

FIG. 2 illustrates a cross-section of a portion of a gas turbine engine.

FIG. 3 illustrates a component that can be incorporated into a gasturbine engine.

FIG. 4 illustrates another view of the component of FIG. 3.

FIG. 5 illustrates a platform of a component.

FIG. 6 illustrates another platform of a component.

FIG. 7 illustrates augmentation features that can be incorporated into acomponent of a gas turbine engine.

FIG. 8 illustrates additional augmentation features that can beincorporated into a component of a gas turbine engine.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The exemplarygas turbine engine 20 is a two-spool turbofan engine that generallyincorporates a fan section 22, a compressor section 24, a combustorsection 26 and a turbine section 28. Alternative engines might includean augmenter section (not shown) among other systems for features. Thefan section 22 drives air along a bypass flow path B, while thecompressor section 24 drives air along a core flow path C forcompression and communication into the combustor section 26. The hotcombustion gases generated in the combustor section 26 are expandedthrough the turbine section 28. Although depicted as a turbofan gasturbine engine in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited toturbofan engines and these teachings could extend to other types ofengines, including but not limited to, three-spool engine architectures.

The gas turbine engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centerlinelongitudinal axis A. The low speed spool 30 and the high speed spool 32may be mounted relative to an engine static structure 33 via severalbearing systems 31. It should be understood that other bearing systems31 may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 34 thatinterconnects a fan 36, a low pressure compressor 38 and a low pressureturbine 39. The inner shaft 34 can be connected to the fan 36 through ageared architecture 45 to drive the fan 36 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 35 thatinterconnects a high pressure compressor 37 and a high pressure turbine40. In this embodiment, the inner shaft 34 and the outer shaft 35 aresupported at various axial locations by bearing systems 31 positionedwithin the engine static structure 33.

A combustor 42 is arranged between the high pressure compressor 37 andthe high pressure turbine 40. A mid-turbine frame 44 may be arrangedgenerally between the high pressure turbine 40 and the low pressureturbine 39. The mid-turbine frame 44 can support one or more bearingsystems 31 of the turbine section 28. The mid-turbine frame 44 mayinclude one or more airfoils 46 that extend within the core flow path C.

The inner shaft 34 and the outer shaft 35 are concentric and rotate viathe bearing systems 31 about the engine centerline longitudinal axis A,which is co-linear with their longitudinal axes. The core airflow iscompressed by the low pressure compressor 38 and the high pressurecompressor 37, is mixed with fuel and burned in the combustor 42, and isthen expanded over the high pressure turbine 40 and the low pressureturbine 39. The high pressure turbine 40 and the low pressure turbine 39rotationally drive the respective high speed spool 32 and the low speedspool 30 in response to the expansion.

Each of the compressor section 24 and the turbine section 28 may includealternating rows of rotor assemblies and vane assemblies (shownschematically) that carry airfoils that extend into the core flow pathC. For example, the rotor assemblies can carry a plurality of rotatingblades 25, while each vane assembly can carry a plurality of vanes 27that extend into the core flow path C. The blades 25 of the rotorassemblies create or extract energy (in the form of pressure) from thecore airflow that is communicated through the gas turbine engine 20along the core flow path C. The vanes 27 of the vane assemblies directthe core air flow to the blades 25 to either add or extract energy.

Various components of the gas turbine engine 20, such as the blades 25and the vanes 27 on the compressor section 24 and/or the turbine section28, may be subjected to repetitive thermal cycling under widely rangingtemperatures and pressures. The hardware of the turbine section 28 isparticularly subjected to relatively extreme operating conditions.Therefore, some components may need cooled during engine operation.Example cooling features that can be incorporated into the components toimprove cooling efficiency are described below.

FIG. 2 schematically illustrates a portion 100 of a gas turbine engine,such as the gas turbine engine 20 of FIG. 1. In this exemplaryembodiment, the portion 100 represents part of the turbine section 28 ofthe gas turbine engine 20. However, it should be understood that otherportions of the gas turbine engine 20 could benefit from the teachingsof this disclosure, including but not limited to, the compressor section24.

In this exemplary embodiment, the portion 100 includes a first component50 and a second component 52 positioned adjacent to the first component50. For example, the first component 50 may represent a vane (agenerally static structure) and the second component 52 may represent ablade mounted for rotation about the engine centerline longitudinal axisA. Although only a single vane and a single blade are illustrated inFIG. 2, multiple vanes and blades could be circumferentially disposedabout the engine centerline longitudinal axis A to provide vane androtor assemblies. The portion 100 could also include additional,alternating rows of vanes and blades. FIG. 2 is highly schematic, and itshould be understood that the various features depicted by this figureare not necessarily drawn to the scale they would be in practice.

In this embodiment, the first component 50 establishes a radially outerand radially inner flow path boundary of the core flow path C anddirects the hot combustion gases communicated along the core flow path Cto the second component 52. The second component 52 rotates to extractenergy from the hot combustion gases that are communicated through thegas turbine engine 20.

The first component 50 and the second component 52 are mounted withinthe portion 100 such that a gap 56 extends between the first component50 and the second component 52. A positive pressure can be maintainedwithin the portion 100 by communicating a leakage airflow 60 into thegap 56. The leakage airflow 60 is communicated through a cavity 58(positioned radially inwardly from the first component 50 and the secondcomponent 52) and then through the gap 56 to keep the hot combustiongases of the core flow path C from entering through the gap 56 andpotentially damaging components. The leakage airflow 60 may becommunicated from the compressor section 24 or some other upstreamlocation of the gas turbine engine 20.

As is discussed in greater detail below, the leakage airflow 60 can alsobe used to cool portions of one or both of the first component 50 andthe second component 52. In other words, the leakage airflow 60 that maybe used to cool portions of the first component 50 and/or the secondcomponent 52 is not a dedicated cooling airflow that serves no otherpurpose other than to cool the component(s) 50, 52. In this disclosure,the term “dedicated cooling airflow” may refer to air which feeds theinside of the component(s) 50, 52 and the term “leakage airflow” mayrefer to airflow that bypasses the inside of the component(s) 50, 52,such as for purging cavities or preventing ingestion.

The first component 50 and/or the second component 52 can includecooling features that increase a local heat transfer effect of the firstand/or second component 50, 52 without requiring a large flow pressureratio. For example, in one embodiment, the first component 50 includes aplatform 64A and the second component 52 includes a platform 64B. Eachof the platforms 64A, 64B includes an outer surface 74 and an innersurface 76. In one embodiment, at least a portion of the secondcomponent 52 extends radially inward from, or under, the first component50.

The platform 64A of the first component may include one or moreaugmentation features 78 disposed on the outer surface 74 for increasingthe heat transfer effect between the platform 64A and the leakageairflow 60. The platform 64A can be cooled by circulating the leakageairflow 60 over the augmentation features 78. In one embodiment, theaugmentation features 78 are disposed on a portion 55 of the platform64A of the first component 50 that axially overlaps the platform 64B ofthe second component 52. The platform 64B could also include one or moreaugmentation features on its outer surface 74, although not shown inthis embodiment.

FIGS. 3 and 4 illustrate a component 150 that can be incorporated into agas turbine engine, such as the gas turbine engine 20 of FIG. 1. In thisembodiment, the component 150 is a turbine vane similar to the firstcomponent 50 of FIG. 2. However, the various features described hereinwith respect to the component 150 could extend to other components ofthe gas turbine engine 20, including but not limited to blades (i.e.,the second component 52 of FIG. 2).

The component 150 of this embodiment includes an inner diameter platform62, an outer diameter platform 64 and an airfoil 66 that extends betweenthe inner diameter platform 62 and the outer diameter platform 64. Eachof the inner diameter platform 62 and the outer diameter platform 64includes a leading edge portion 68, a trailing edge portion 70 andopposing mate faces 71, 73. The inner diameter and outer diameterplatforms 62, 64 axially extend between the leading edge portion 68 andthe trailing edge portion 70 and circumferentially extend between theopposing mate faces 71, 73. The opposing mate faces 71, 73 can bemounted relative to corresponding mate faces of adjacent components of agas turbine engine to provide a full ring assembly, such as a full ringvane assembly that can be circumferentially disposed about the enginecenterline longitudinal axis A (see FIG. 1).

The inner diameter and outer diameter platforms 62, 64 can also includean outer surface 74 (for example, a non-gas path side) and an innersurface 76 (a gas path side). In other words, when the component 150 ismounted within a gas turbine engine 20, the outer surfaces 74 arepositioned on a non-core flow path side of the component 150, while theinner surfaces 76 establish the outer boundaries of the core flow path Cof the gas turbine engine 20.

One or both of the inner diameter platform 62 and the outer diameterplatform 64 can include one or more augmentation features 78 disposed onthe outer surfaces 74 of the inner diameter platform 62 and/or the outerdiameter platform 64. In this embodiment, the augmentation features 78are positioned at the trailing edge portion 70 of the inner diameterplatform 62 (see FIG. 2 and FIG. 4). In another embodiment, theaugmentation features 78 may be disposed at the leading edge portion 68of the inner diameter platform 62 and/or the outer diameter platform 64(see FIG. 5). In yet another embodiment, the augmentation features 78may be formed on both the leading edge portion 68 and the trailing edgeportion 70 of the inner diameter platform 62 and/or the outer diameterplatform 64 (see FIG. 6). The augmentation features 78 may be disposedon any portion of the inner diameter platform 62 and/or the outerdiameter platform 64, including any portion of the inner or outerdiameter platforms 62, 64 that axially overlap a neighboring component(see portion 55 of FIG. 2, for example).

The exemplary augmentation features 78 can include any heat transferaugmentation features including but not limited to trip strips, pinfins, chevron trip strips, or any combination of features. In thisembodiment, the augmentation features 78 include trip strips. Theaugmentation features 78 turbulate the flow of the leakage airflow 60that comes into contact with the component 150 (as shown in FIG. 2) andadds surface area to platform(s) 62, 64 to enhance heat transfer betweenthe leakage airflow 60 and the platform(s) 62, 64.

FIG. 7 illustrates additional features of the augmentation features 78described above. The illustrated platform 62, 64 could be representativeof either an inner diameter platform or an outer diameter platform. Inthis embodiment, the augmentation features 78 include a plurality oftrip strips 78A that are circumferentially spaced between the opposingmate faces 71, 73 at a trailing edge portion 70 of the platform 62, 64.The trip strips 78A extend parallel to one another and may each beangled at an angle α relative to the opposing mate faces 71, 73. Thevalue of the angle α can vary depending on design specific criteria,including but not limited to the amount of heat transfer required tocool the platform 62, 64. The leakage airflow 60 may be circulated overthe trip strips 78A in both a circumferential direction CD as well as anaxial direction AD to cool the platform 62, 64.

FIG. 8 illustrates another plurality of augmentation features 178 thatcan be incorporated into a platform 62, 64 of a component 150. In thisembodiment, the augmentation features 178 include a plurality of tripstrips 178A that are circumferentially spaced between the opposing matefaces 71, 73 at a trailing edge portion 70 of the platform 62, 64. Theplurality of trip strips 178A of this embodiment include a first portion80 and a second portion 82 that is transverse to the first portion 80.The first portions 80 of the plurality of trip strips 178A extendparallel to one another and may each be angled at an angle α relative tothe opposing mate faces 71, 73. The second portions 82 of the pluralityof trip strips 178A also extend parallel to one another and may beangled at an angle β relative to the opposing mate faces 71, 73. In thisembodiment, the angle β is a different angle than the angle α. Thesecond portions 82 of the plurality of trip strips 178A direct theleakage airflow 60 into the core flow path C by directing the leakageairflow 60 circumferentially along the direction of rotation of aneighboring component (shown schematically via arrow 110), such as ablade (see the second component 52 of FIG. 2), thereby providingimproved heat transfer and reducing mixing loss.

Although the different non-limiting embodiments are illustrated ashaving specific components, the embodiments of this disclosure are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the non-limitingembodiments.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed and illustrated in these exemplary embodiments,other arrangements could also benefit from the teachings of thisdisclosure.

The foregoing description shall be interpreted as illustrative and notin any limiting sense. A worker of ordinary skill in the art wouldrecognize that various modifications could come within the scope of thisdisclosure. For these reasons, the following claims should be studied todetermine the true scope and content of this disclosure.

What is claimed is:
 1. A component for a gas turbine engine, comprising:a platform having an outer surface and an inner surface that axiallyextend between a leading edge portion and a trailing edge portion; andat least one augmentation feature disposed on at least one of saidleading edge portion and said trailing edge portion of said outersurface of said platform.
 2. The component as recited in claim 1,wherein said platform is an inner diameter platform.
 3. The component asrecited in claim 1, wherein the component is a turbine vane.
 4. Thecomponent as recited in claim 1, wherein said outer surface is a non-gaspath side of said platform and said inner surface is a gas path side ofsaid platform.
 5. The component as recited in claim 1, comprising anairfoil that extends from said inner surface of said platform.
 6. Thecomponent as recited in claim 1, wherein said at least one augmentationfeature includes a plurality of trip strips.
 7. The component as recitedin claim 6, wherein said plurality of trip strips are disposed at saidtrailing edge portion of said platform.
 8. The component as recited inclaim 6, wherein each of said plurality of trip strips are angledrelative to opposing mate faces of said platform.
 9. The component asrecited in claim 6, wherein each of said plurality of trip stripsinclude a first portion and a second portion that is transverse to thefirst portion.
 10. The component as recited in claim 9, wherein saidfirst portions are angled at a first angle relative to a mate face ofsaid platform and said second portions are angled at a second angledifferent from said first angle relative to said mate face.
 11. Thecomponent as recited in claim 1, wherein said at least one augmentationfeature is disposed at said leading edge portion of said platform. 12.The component as recited in claim 1, wherein said at least oneaugmentation feature is disposed on a portion of said outer surface ofsaid platform that axially overlaps a neighboring component of the gasturbine engine.
 13. A gas turbine engine, comprising: a compressorsection; a combustor section in fluid communication with said compressorsection; a turbine section in fluid communication with said combustorsection; and wherein at least one of said compressor section and saidturbine section includes: a first component having a platform thatincludes an outer surface and an inner surface that axially extendbetween a leading edge portion and a trailing edge portion; a secondcomponent mounted adjacent to said first component and including aplatform; wherein a portion of said platform of said first componentaxially overlaps a portion of said platform of said second component,and said portion of said first platform includes at least oneaugmentation feature disposed on said outer surface of said platform.14. The gas turbine engine as recited in claim 13, wherein said firstcomponent is a vane and said second component is a blade.
 15. The gasturbine engine as recited in claim 13, wherein said at least oneaugmentation feature includes a plurality of trip strips.
 16. The gasturbine engine as recited in claim 13, wherein said at least oneaugmentation feature is disposed on at least one of said leading edgeportion and said trailing edge portion of said platform of said firstcomponent.
 17. The gas turbine engine as recited in claim 13, whereinsaid at least one augmentation feature is angled relative to a mate faceof said platform.
 18. A method of cooling a component of a gas turbineengine, comprising the steps of: cooling a platform of the componentwith a leakage airflow that is communicated from a cavity positionedradially inwardly from the component by circulating the leakage airflowover at least one augmentation feature that is disposed on an outersurface of the platform.
 19. The method as recited in claim 18, whereinthe leakage airflow is not a dedicated cooling airflow that iscommunicated inside of the component.
 20. The method as recited in claim18, wherein the at least one augmentation feature is disposed on atrailing edge portion of the platform.